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NACA 0010-34 (naca001034-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: NACA 0010-34 (naca001034-il)
Reynolds number: 200,000
Max Cl/Cd: 38.57 at α=3°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-naca001034-il-200000-n5.txt
Download as CSV file: xf-naca001034-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 0010-34                                    
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.6329   0.09017   0.08675  -0.0204   1.0000   0.0100
 -10.000  -0.6425   0.08180   0.07838  -0.0271   1.0000   0.0095
  -9.750  -0.6575   0.07483   0.07135  -0.0324   1.0000   0.0092
  -9.500  -0.6746   0.06950   0.06593  -0.0348   1.0000   0.0089
  -9.250  -0.6919   0.06523   0.06156  -0.0347   1.0000   0.0087
  -9.000  -0.7082   0.06172   0.05792  -0.0326   1.0000   0.0086
  -8.750  -0.7193   0.05814   0.05417  -0.0302   1.0000   0.0087
  -8.500  -0.7261   0.05460   0.05044  -0.0276   1.0000   0.0088
  -8.250  -0.7296   0.05122   0.04684  -0.0248   1.0000   0.0091
  -8.000  -0.7306   0.04798   0.04335  -0.0216   1.0000   0.0094
  -7.750  -0.7304   0.04482   0.03992  -0.0181   1.0000   0.0096
  -7.500  -0.7291   0.04165   0.03645  -0.0145   1.0000   0.0096
  -7.250  -0.7264   0.03845   0.03293  -0.0108   1.0000   0.0095
  -7.000  -0.7211   0.03541   0.02954  -0.0072   1.0000   0.0093
  -6.750  -0.7131   0.03256   0.02628  -0.0037   1.0000   0.0093
  -6.500  -0.7022   0.02994   0.02328  -0.0006   1.0000   0.0092
  -6.250  -0.6885   0.02757   0.02055   0.0022   1.0000   0.0093
  -6.000  -0.6723   0.02546   0.01811   0.0045   1.0000   0.0095
  -5.750  -0.6542   0.02361   0.01597   0.0066   1.0000   0.0098
  -5.500  -0.6351   0.02208   0.01422   0.0084   1.0000   0.0103
  -5.250  -0.6158   0.02145   0.01343   0.0100   1.0000   0.0113
  -5.000  -0.5986   0.01911   0.01093   0.0121   1.0000   0.0125
  -4.750  -0.5724   0.01771   0.00944   0.0121   0.9979   0.0135
  -4.500  -0.5432   0.01664   0.00823   0.0114   0.9952   0.0151
  -4.250  -0.5136   0.01589   0.00737   0.0107   0.9919   0.0184
  -4.000  -0.4833   0.01511   0.00641   0.0098   0.9886   0.0217
  -3.750  -0.4523   0.01441   0.00565   0.0088   0.9856   0.0330
  -3.500  -0.4268   0.01319   0.00491   0.0085   0.9812   0.1335
  -3.250  -0.4060   0.01130   0.00434   0.0084   0.9776   0.4241
  -3.000  -0.3892   0.01010   0.00428   0.0106   0.9723   0.6597
  -2.750  -0.3612   0.00988   0.00437   0.0110   0.9687   0.7559
  -2.500  -0.3276   0.00986   0.00435   0.0101   0.9661   0.7949
  -2.250  -0.2983   0.00985   0.00427   0.0098   0.9607   0.8154
  -2.000  -0.2649   0.00986   0.00422   0.0087   0.9569   0.8326
  -1.750  -0.2292   0.00988   0.00415   0.0071   0.9540   0.8470
  -1.500  -0.1984   0.00989   0.00411   0.0065   0.9486   0.8598
  -1.250  -0.1660   0.00989   0.00407   0.0055   0.9435   0.8714
  -1.000  -0.1301   0.00990   0.00402   0.0038   0.9399   0.8808
  -0.750  -0.0991   0.00992   0.00401   0.0032   0.9332   0.8898
  -0.500  -0.0660   0.00990   0.00397   0.0021   0.9276   0.8989
  -0.250  -0.0327   0.00992   0.00397   0.0010   0.9209   0.9061
   0.000   0.0000   0.00991   0.00395   0.0000   0.9140   0.9140
   0.250   0.0328   0.00992   0.00397  -0.0010   0.9061   0.9209
   0.500   0.0661   0.00990   0.00397  -0.0021   0.8989   0.9276
   0.750   0.0992   0.00992   0.00401  -0.0032   0.8898   0.9332
   1.000   0.1302   0.00990   0.00402  -0.0038   0.8808   0.9399
   1.250   0.1661   0.00989   0.00406  -0.0055   0.8714   0.9435
   1.500   0.1984   0.00989   0.00411  -0.0065   0.8598   0.9487
   1.750   0.2293   0.00987   0.00415  -0.0071   0.8470   0.9540
   2.000   0.2649   0.00986   0.00422  -0.0087   0.8326   0.9570
   2.250   0.2984   0.00985   0.00426  -0.0098   0.8154   0.9608
   2.500   0.3277   0.00986   0.00435  -0.0101   0.7950   0.9661
   2.750   0.3612   0.00987   0.00437  -0.0110   0.7558   0.9687
   3.000   0.3892   0.01009   0.00427  -0.0106   0.6601   0.9723
   3.250   0.4060   0.01130   0.00433  -0.0084   0.4241   0.9776
   3.750   0.4523   0.01442   0.00565  -0.0088   0.0329   0.9857
   4.000   0.4834   0.01510   0.00641  -0.0099   0.0218   0.9886
   4.250   0.5137   0.01589   0.00737  -0.0107   0.0185   0.9919
   4.500   0.5434   0.01664   0.00823  -0.0115   0.0151   0.9953
   4.750   0.5726   0.01771   0.00944  -0.0121   0.0135   0.9979
   5.000   0.5984   0.01911   0.01093  -0.0120   0.0125   1.0000
   5.250   0.6157   0.02149   0.01347  -0.0100   0.0112   1.0000
   5.500   0.6350   0.02209   0.01422  -0.0084   0.0103   1.0000
   5.750   0.6541   0.02360   0.01596  -0.0066   0.0098   1.0000
   6.000   0.6722   0.02545   0.01810  -0.0045   0.0095   1.0000
   6.250   0.6884   0.02756   0.02054  -0.0021   0.0093   1.0000
   6.500   0.7021   0.02993   0.02327   0.0006   0.0092   1.0000
   6.750   0.7130   0.03256   0.02627   0.0038   0.0093   1.0000
   7.000   0.7210   0.03541   0.02954   0.0072   0.0093   1.0000
   7.250   0.7263   0.03845   0.03293   0.0108   0.0095   1.0000
   7.500   0.7291   0.04165   0.03645   0.0145   0.0096   1.0000
   7.750   0.7304   0.04482   0.03992   0.0181   0.0096   1.0000
   8.000   0.7309   0.04797   0.04334   0.0216   0.0094   1.0000
   8.250   0.7298   0.05121   0.04683   0.0247   0.0091   1.0000
   8.500   0.7265   0.05458   0.05042   0.0276   0.0088   1.0000
   8.750   0.7196   0.05814   0.05417   0.0301   0.0086   1.0000
   9.000   0.7084   0.06176   0.05796   0.0325   0.0087   1.0000
   9.250   0.6922   0.06529   0.06162   0.0346   0.0087   1.0000
   9.500   0.6751   0.06955   0.06599   0.0347   0.0089   1.0000
   9.750   0.6582   0.07491   0.07143   0.0322   0.0091   1.0000
  10.000   0.6432   0.08194   0.07853   0.0268   0.0095   1.0000
  10.250   0.6342   0.09034   0.08692   0.0200   0.0099   1.0000
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